This present invention relates generally to flight control actuation systems and, more specifically, to a method and apparatus for a dual actuator control system, containing at least one electromechanical actuator and at least one pneumatic actuator. The present invention concerns actuator systems for controlling flight control surfaces on aircraft, spacecraft, missiles, and other flight vehicles.
Actuator servomechanism systems are used to manipulate flight control surfaces to control flight direction, speed, inclination and other positional adjustments for flight vehicles. The actuator systems have used mechanical, hydraulic, electrical, piezeoelectrical, and electromechanical systems to apply force to the control surfaces. For safety, redundant parallel systems are used to independently maintain control of the flight control surface in the event of failure of one of the actuator systems. One such parallel system is disclosed in U.S. Pat. No. 5,074,495 to Raymond. The hydraulically- and electrically-powered actuators individually are capable of providing full actuation power. This system design does not account for significant variances from the normal operational range of the electrically powered actuator, such as control surface flutter and shockwave conditions. Flutter is oscillatory motion between the vehicle frame and the control surface. Flutter increases as the vehicle approaches resonant frequencies. Shockwave conditions increase control surface loads as the vehicle approaches sonic velocity. To account for the resultant high control surface loads, the actuator systems must be large in size and mass, negatively impacting flight vehicle weight constraints and aerodynamic envelope limitations. Additionally, large flight vehicles traveling at high speeds introduce risks of overloading the electrical actuator from the greater forces needed to manipulate the flight control surfaces in such situations. To address these issues, power-assist systems were developed to amplify the force applied from the main control system and to minimize the control system resistance to movement. An example of such a system is disclosed in U.S. Pat. No. 6,349,900 to Uttley, et al. This actuator system uses an electrical actuator assisted by a control tab mounted on the control surface. This system's drawbacks are lower output forces than conventional actuator systems, and the excess size and mass added to the flight vehicle from the use of control tabs.
None of the prior art is specifically intended for lightweight, high-speed applications, and some suffer from one or more of the following disadvantages:                a) excessive mass and size.        b) inability to accommodate flutter or shockwave effects.        c) increased cooling requirements.        d) low achievable output forces.        e) inferior aerodynamic envelope conditions.        f) inability to use detected electrical actuator current variations.        
As can be seen, there is a need for an improved apparatus and method for a light, small, amplified flight control actuation system, which reacts well to flight extremes, such as high speeds and resonant frequencies, does not require excessive cooling, provides high output forces and adapts to detected electrical actuator current variations.